Gas-Generator Augmented Expander Cycle Rocket Engine

ABSTRACT

An augmented expander cycle rocket engine includes first and second turbopumps for respectively pumping fuel and oxidizer. A gas-generator receives a first portion of fuel output from the first turbopump and a first portion of oxidizer output from the second turbopump to ignite and discharge heated gas. A heat exchanger close-coupled to the gas-generator receives in a first conduit the discharged heated gas, and transfers heat to an adjacent second conduit carrying fuel exiting the cooling passages of a primary combustion chamber. Heat is transferred to the fuel passing through the cooling passages. The heated fuel enters the second conduit of the heat exchanger to absorb more heat from the first conduit, and then flows to drive a turbine of one or both of the turbopumps. The arrangement prevents the turbopumps exposure to combusted gas that could freeze in the turbomachinery and cause catastrophic failure upon attempted engine restart.

ORIGIN OF THE INVENTION

This invention was made by an employee of the United States Governmentand may be manufactured and used by or for the Government forgovernmental purposes without the payment of any royalties.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The present invention relates to rocket engines. More particularly, theinvention relates to upper stage, restart capable, highly reliablerocket engines.

2. Discussion of the Related Art

Launch vehicles are typically comprised of multiple elements stackedtogether in stages. There are lower stages and upper stages, with thelower stages being used for lifting the most weight and pushing avehicle through the thicker layers of lower atmosphere. The upper stagesnormally fire subsequent to the vehicle traveling very fast in thinnerlayers of atmosphere and at a very high rate of speed. Accordingly,improvements in upper stage rocket design, even relatively minorimprovement can provide a large increase in the performance of amission.

For example, an improvement in performance for an upper stage canpositively affect the amount of payload that a rocket can place intoorbit. As a result, some of the most advanced rocket research hasfocused on upper stages. Although there is no strict definition of an“upper stage,” it usually refers to the second, third, and fourth (ifany) stages of a rocket, ignited and fired at high altitude.

Space missions require rocket engines that provide high thrust, highefficiency, and robust durability in order to be operated under thedemanding conditions of outer space. For example, there are extremedifferences in temperature and pressure both during takeoff and inflight.

There have been many attempts to develop high performance and highreliability rocket engines. One such engine is the standardgas-generator cycle, wherein a small fraction of the overall inletpropellant flow is combusted in a secondary combustion chamber, and theproduct of this combustion is used to drive the turbomachinery directly.These combustion gases are then discharged. Some functioning or pastexamples of this engine cycle include the J-2 (Saturn V) and the lowerstage engine RS-68 (Delta IV).

A concept closely related but more complex than the gas-generator cycleis called the staged-combustion cycle. In this case rather thandischarging the combustion gases from the secondary combustion zone,these gases are introduced into the primary combustion chamber andre-combusted. The result is an engine with extremely high performancebut also a high degree of complexity. A functioning example of this typeof engine is the main engine of the NASA Space Shuttle.

Yet another typical rocket engine cycle proposed for upper stageapplication is the tap-off cycle, wherein hot combustion products aretapped off from the primary combustion zone and these gases are used todrive the turbomachinery before being discharged, as in thegas-generator cycle. A functioning example of this type of engine, butone that was never actually realized on a rocket vehicle stage is theJ-2S developed by the U.S. during the early 1970's.

Finally, the expander cycle rocket engine, which is well known in theprior art, presents an ideal choice for upper stage use. In general, thefuel is heated before it is combusted, typically with waste heat fromthe main combustion chamber. The expander cycle is based upon theconcept of driving the turbomachinery with gases warmed throughregenerative cooling of the thrust chamber assembly, so as to eliminatethe need for a secondary combustion zone.

The expander cycle rocket engine has many inherent benefits over othertypical cycles such as the standard gas-generator cycle, the tap-offcycle, or the staged combustion cycle. In an expander cycle, the fuel istypically heated before it is combusted, the heat being supplied bywaste heat from a main combustion chamber.

FIG. 1 is a simplified illustration of a typical expander engine (closedcycle). The heat from the nozzle 105 and the combustion chamber 101 areused to power a fuel pump 115 and an oxidizer pump 120. As the liquidfuel passes through coolant passages 125 in the walls of the combustionchamber 101, the fuel, typically at supercritical pressures, picks upenergy in the form of heat. This energy increase within the coolantpassages 125 of the walls of the combustion chamber is sufficient todrive the gas turbine 130. The gas turbine provides the power necessaryto drive the fuel pump 115 and the oxidizer pump 120. The turbinedischarge along with the pumped oxidizer are then provided to thecombustion chamber 101 for combustion.

In a closed-cycle expander engine, the exhaust is sent from the turbineto the combustion chamber, whereas in an open-cycle expander engine onlysome of the fuel is heated to power the turbines, and then vented,resulting in decreased overall efficiency, though this design does haveother potential benefits.

Some disadvantages of all of the aforementioned expander cycle enginesincludes a limit in the amount of available power to drive theturbomachinery because the driving of the turbomachinery is caused byusing the heat extracted through the cooling wall of the primarycombustion chamber. Attempting to increase the amount of heattransferred can normally be achieved by using extremely thin wallsand/or exotic means of increasing the local wall temperature within thecombustion chamber.

However, to design a thinner wall or increase the wall temperaturecauses a reduction of the structural strength of the combustion chamberwall material, thereby reducing the reliability of the component and theengine. Also, typically such engines are shutdown fuel-rich, meaningthat the through cooling passages permit cryogenic hydrogen to flowduring the shutdown process. Shutting down the engine fuel-rich isstandard practice and is intended to establish a benign environment andto avoid a catastrophic failure. However, when this practice isperformed in outer space, the hardware and the passages can stay verycold for a long time. Due to the fact that the expander cycle requiresheat to initiate and drive the cycle, restarting the engine in this coldstate is difficult and unreliable, as it will take some time before theengine has become warm enough to initiate the start sequence.

FIGS. 2 and 3 show an example of one attempt at improving the expandercycle engine as disclosed in U.S. Pat. No. 6,832,471 to Hewitt. Hewittdiscloses that by injecting the oxidizer in two streams, with a smallerstream being injected into the upstream or preburner, and the remainderto the downstream or main combustion section, the use of a coolingelement with a high intimate heat exchange construction is permitted toextract a high level of energy from the preburner gas in the form ofheat without damaging the cooling element.

More specifically, Hewitt discloses a nozzle 11 of an expander cyclesupersonic rocket engine. The drawing shows a combustion chamber 13, athroat 17, a supersonic section or skirt 15. The combustion chamber 13has an upstream or preburner section 21 and a main downstream section22. The upstream or preburner section 21 is a secondary combustion zone,which is used instead of the more common gas-generator and a maindownstream section. In this configuration, the products from thesecondary combustion zone are directly fed directly into the primarycombustion zone rather than discharged externally.

Still referring to FIGS. 2 and 3, Hewitt discloses that the firstportion of liquid oxygen 23 is fed to an inlet torus 24 that surroundsthe upstream portion, wherein the torus directs the liquid oxygenthrough the chamber wall and into the interior of the chamber. Thesecond portion (remainder) of the liquid oxygen is fed to inlet torus 25and the heated gaseous hydrogen 27 from the turbopump turbine 41 (FIG.3) exhaust is fed to the preburner section 21 for combustion. A plateletlaminate 31 consisting of a laminate of two stacks of circular disks33,34, one above the other, separated by a barrier disk 35, is used forheat exchange. The circular disks 33, 34 have central openings 32 and anopen space 36 at the stack periphery or by axial channels to form twoindependent flowpaths that are in heat exchange relationship but notfluid communication. One flowpath is for the combustion gas, and theother is for the uncombusted hydrogen fuel that serves as the coolant.

Still referring to FIGS. 2 and 3, Hewitt discloses the flowpath forcombustion gases passes radially outward through the upstream stack 33,then into the annular space 36, then radially inward through thedownstream stack 34, then through tubular passages in a distributionmanifold 37. The flowpath for the hydrogen fuel acting as a coolantenters the downstream stack 34 upon emerging from the jacket 16, thenflows radially inward through the downstream stack 34 (counter-currentto the combustion gas) then through axial passages that connect thedownstream stack 34 to upstream stack 33, then radially outward throughthe upstream stack 33 to a space above the upstream stack that leadsoutwards 38.

The coolant 38 being led outward is now in a gaseous form and directedto the drive turbine 41 of the turbopump (shown in FIG. 3). The heatedgas 40 drives turbine 41, which drives two shafts 42, 23 and hasseparate pumps 44, 45 for liquid hydrogen and fuel. The heated gas 40pumps fresh coolant to the jacket 16, which flows in the directionindicated by dashed arrows 17,18 while the heated gas itself is expandedand fed 27 to the preburner injector for combustion in the preburner andthe main section of the combustion chamber. The partially cooledcombustion gas from the preburner 21 is joined by the remainder of theliquid oxygen feed at the downstream face of the injector/manifold 37 todistribute both the fuel-rich preburner gas and freshly supplied oxygenacross the width of the combustion chamber. The expanded uncombustedhydrogen 46 that emerges from the drive turbine 41 is then injected intothe combustion side of the upstream section 21 of the engine.

Thus, in Hewitt the preburner combustion gas is cooled in asubstantially uniform manner to a moderate temperature by cooling thebulk of the gas rather than cooling only the gas in a boundary layeradjacent to the chamber wall.

However Hewitt has disadvantages because the primary and secondarycombustion zones are so closely linked. Due to the fact that Hewitteffectively incorporates a staged-combustion element within the expandercycle, this inherently brings into the situation the difficult balancingact of the interplay between the two combustion zones. Hewitt alsosuffers from some of the disadvantages of the other engines as well.

Other drawbacks of any of the above-mentioned rocket engines, includingHewitt, include the problem that after the engine is shut down, theresidual combustion products, specifically steam, can freeze in theinjectors and cause damage. Moreover, for those engine cycles usingcombustion products as the turbine drive gas, the steam can also freezein the turbomachinery, which has the potential to be even moredisastrous upon attempted engine restart than the potential damage fromsteam freezing in the injectors. Thus, there is a need in the art for animproved expander cycle rocket engine.

SUMMARY OF THE INVENTION

The present invention provides an improvement over the standard expandercycle engine by including a gas-generator and a close-coupled heatexchanger. The present invention also provides a significant improvementover the pre-burner system as disclosed in Hewitt because by completelydecoupling two combustion zones, there is no delicate balance ofperformance between the two combustion zones. Due to the fact that inthe present invention the gas-generator is not directly coupled to theturbomachinery, the possibility that combustion products, specificallysteam, can get trapped in the turbomachinery components and preclude thepossibility of ice forming in the turbomachinery prior to a subsequentrestart of the engine, (which is a potentially catastrophic event) ispractically eliminated.

According to a first aspect of the invention, there are two turbopumps,one for liquid hydrogen (for use as a fuel) and one for liquid oxygen(for use as an oxidizer). There is also a thrust chamber assembly withinwhich the primary combustion takes place in order to provide thepropulsive discharge of gases. The thrust chamber assembly is made in amanner allowing for fuel to flow interior to the wall, either via tubesor milled channels, thereby providing cooling.

Compared with conventional expander cycle rocket engines, the rocketengine according to the present invention includes a gas-generatorcomprising a relatively small combustion chamber that is fed propellants(fuel and oxidizer), directly off the two pump discharges. Flow to thecombustion chamber is controlled by two valves, one for fuel and one forthe oxidizer.

According to the invention, the gas-generator is ignited first, prior tothe main combustion chamber, immediately causing heat to be imparted tothe coolant hydrogen. The result is that power is more quickly andassuredly delivered to the two turbopumps, thereby ensuring a smoothengine start transient. During regular operation, the gas-generator canbe modulated to provide for various levels of engine thrust and can evenbe extinguished so the engine can function like a typical expander cycleif so desired. For example, a control module (not shown) provides, upona start or restart, an actuation sequence in which the gas-generatorfuel valve and the gas-generator oxidizer valve are actuated open priorto opening the coolant control valve, main fuel valve and main oxidizervalve for igniting the gas-generator prior to igniting the primarycombustion chamber. This ignition sequence ensures a smooth engine starttransient

According to the invention, the turbines of the turbopumps are drivensoley by the heated fuel and no combusted gas flows from either thegas-generator or the primary combustion chamber to directly drive theturbines.

In addition, the invention may include a secondary, smaller nozzle todischarge the gas-generator gases, instead of discharging them thoughthe aft end of the first nozzle. The use of a secondary nozzle canprovide the function of a settling motor, by using the ignition, initialburning and discharge from the gas-generator as the stage settlingthrust, as a small amount of initial thrust is needed to locate thepropellants in the proper manner when the engine is started in thezero-acceleration state.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a schematic of one type of known expander cycle rocketengine.

FIGS. 2 and 3 show the nozzle of an expander cycle and a simplifieddrawing of an expander cycle rocket engine as disclosed by Hewitt inU.S. Pat. No. 6,832,471.

FIG. 4A shows one embodiment of a flow circuit for an expander cyclerocket engine according to the present invention.

FIG. 4B is variation of the embodiment of the flow circuit shown in FIG.4A.

DETAILED DESCRIPTION OF THE INVENTION

The descriptions provided herein in accordance with the drawings havebeen provided for purposes of illustration and not for limitation. Aperson of ordinary skill in the art understands and appreciates thatthere are many variations of the present invention not shown that do notdepart from the spirit of the invention and the scope of the appendedclaims.

FIG. 4A is a simplified view of a first embodiment of a gas-generatoraugmented rocket engine according to the present invention. A person orordinary skill in the art understands that the flow circuit shown inFIG. 4A is simplified so as not to obscure the invention withunnecessary detail. There are also a number of valves, ancillary lines,and by-pass pathways most of which are not shown on the diagram, butcould typically be included. FIG. 4B uses the same reference numeralsexcept they are increased by 1000, and indicates the same components astheir counterparts in FIG. 4A, except for the second nozzle, which isdifferent.

With reference to FIG. 4A, turbopumps, 201,203, which are a combinationof a pump and a turbine on either end of a common shaft, are used. Afirst turbopump 201 provides liquid hydrogen (fuel), and a secondturbopump 203 provides liquid oxygen (oxidizer), both of which arestored in the vehicle stage and delivered to the engine through twoseparate feedlines 202,204. Through the pumps, the propellants areraised to higher pressure to be able to traverse the rest of the cycle.

The thrust chamber assembly 207 has at its forward or upstream end theprimary combustion chamber 208, within which the primary combustion ofthe fuel and oxidizer takes place to provide the propulsive discharge ofhot gases. At the rearward end or downstream end is a nozzle 207 b, bywhich the gases are discharged to propel the rocket.

The thrust chamber assembly 207 is made in a manner to allow for coolantin the form of fuel to flow interior to the wall, either via tubes ormilled channels, thereby providing cooling to the hot gases in thecombustion chamber. This transfer of heat is also desirable because itis desired to heat the fuel used to drive the turbines.

On the fuel side 202, some propellant 201 a is directed straight to theprimary combustion chamber 208 of thrust assembly 207 through a fueloutput conduit via the main fuel valve 209. However, a fraction 201 b ofthe fuel is flowed through the walls of the combustion chamber viacoolant control valve 211 for the dual purpose of cooling the primarycombustion chamber assembly 208 of thrust chamber assembly 207, and forthe fuel to pick up energy in the form of heat. This same heated fuelthen passes through first conduit 219 a of heat exchanger 219 and isadditionally heated by the gas discharge in second conduit 219 b. Theheated fuel exits and follows path 221 to power the turbines 206 of bothturbopumps 201, 203. When the turbines are driven, the turbo pumps canprovide more fuel and oxidizer to the primary combustion chamber 208.

As an additional element heretofore unknown, another fraction 201 c ofthe fuel pump discharge is conducted to the gas-generator 213. Thegas-generator 213 is a small combustion chamber fed by propellants (theliquid hydrogen and liquid oxygen fuel from turbopumps 201, 203), whichpass through respective gas-generator fuel and oxidizer valves (215,217) that may include actuators to control the type and quantity ofcombustion gas output from the gas-generator 213. Immediately downstreamof the fuel and oxidizer valves is the heat exchanger 219 comprised ofnumerous tubes situated cross-wise to the hot gas flow and allowing fuelto flow within and become heated. For simplicity purposes, FIG. 4A showsa non-limiting example in which first conduit 219 a and second conduit219 b are used for heat exchange.

The fuel following path 201 b is used to cool the thrust chamberassembly 207 (via coolant control valve 211) and then after exiting theprimary combustion chamber 208 of the thrust chamber assembly 207 theheated fuel flows through the first conduit 219 a of the heat exchanger219 situated downstream of the gas-generator 213. In this way, the fuelin path 201 b, which already has received heat while being used to coolthe thrust chamber assembly 207, receives even more energy in the formof heat in the heat exchanger 219 from the heated gas discharge leavingthe gas-generator 213. The fuel, which has now been heated by both thetransfer of heat from primary combustion chamber 208 and the heatexchanger 219, now can power the turbines 206 with more energy andreliability than in rocket engines previously known.

Still referring to FIG. 4A, after passage through the second conduit 219b in the heat exchanger, the discharged gas from the gas-generator 213is typically introduced to the aft end 207 b of the nozzle in the thrustchamber assembly 207.

On the oxidizer side 204 of the flow circuit, the layout is somewhatless complex. Liquid oxygen enters the engine, is pumped by turbopump203 to a higher pressure, and then the majority of the liquid oxygen isthen conducted directly through output conduit path 203 a via valve 212to the primary combustion chamber 208 within the thrust chamber assembly207. A small fraction of the liquid oxygen is tapped off via path 203 bto provide oxidizer for the combustion within the gas-generator 213.

Still referring to FIG. 4A, the use of the gas-generator 213 andclose-coupled heat exchanger 219 permits the heating of the fuel thatpowers the turbines of the turbopumps to a higher temperature than intypical expander cycle rocket engines and other known engines includingthose previously discussed, permitting the turbines to receive atransfer of power faster and at a higher level than known heretofore. Itshould be noted that the turbines 206 are driven only with heated fuel,and not combustion gas, as the combustion gas discharged from gasgenerator 213 travels through second conduit 219 b and then to the aft207 b of the nozzle for discharge. Accordingly, the turbines will not besubjected to steam that would be present in the combustion gas, and thuswhen the engine is turned off, there would not be a possibility of iceforming in the turbomachinery as in several rocket engine cycles knownheretofore.

In operation, the gas-generator 213 would be ignited first causing heatto be immediately imparted to the coolant hydrogen that passes throughthe walls of the combustion chamber of the primary combustion chamberprior to entering the heat exchanger. The result is that power would bedelivered more assuredly and faster to the turbine ends of the twoturbopumps 201, 203, thereby ensuring a smooth engine start transient.

In addition, one advantage of the invention is that additional heattransfer is obtained by decoupling the heat exchanger from the thrustchamber assembly, thereby overcoming the limitations of known rocketengines using regenerative cooling for providing power to the twoturbopumps 201, 203. During the engine start transient, the structure ofthe present invention allows for a more regular, repeatable, andtherefore reliable engine start, even when the thrust chamber assemblyhardware has little latent heat. During regular engine operation, thisdecoupling allows for the thrust chamber walls to be kept at moreregular temperatures, thereby allowing them to maintain their structuralstrength and thus increase engine reliability.

Moreover, another advantage of the present invention over rocket enginesknown heretofore is that an output of the gas-generator does notdirectly drive the turbomachinery. As a result, no combustion productsfrom the gas-generator, such as steam, can get trapped in theturbomachinery components. This invention eliminates the possibility offorming from trapped combustion products in the turbomachinery prior tothe next restart of the engine. Such ice can cause catastrophic results.

A person of ordinary skill in the art also will appreciate andunderstand that the simple injector included as part of thegas-generator according to the present invention would be relativelyeasy to purge of combustion products in comparison to turbomachinery.Also, depending on the operational needs of the rocket engine stage, itmay be possible to run the gas-generator at very low total flow rates oreven extinguish it altogether to achieve a higher level of overallengine performance than is possible with tap-off or gas-generator rocketengine cycles.

During regular operation, the gas-generator 213 can be modulated byvariably actuating the gas-generator fuel valve and/or the gas-generatoroxidizer valve so as to provide various levels of engine thrust, or evenbe extinguished.

One possible variation of the present invention, as shown in FIG. 4B,would be that instead of discharging the gas-generator through the aftend 1207 b of the nozzle 1207, a second smaller nozzle 1207 a can beadded to discharge these gases. Such a secondary nozzle 1207 a has beenused in the past on some gas-generator cycle engines, but not with a gasgenerator augmented expander cycle rocket engine according to thepresent invention. In configuring the engine as shown in FIG. 4B, thesecondary nozzle discharge could function in the capacity of a “settlingmotor”. As the engine is typically started in a zero-acceleration state,in order to locate the propellants in the proper manner within thevehicle stage, a small amount of initial thrust is needed. Thus, theaddition of the secondary nozzle could be used to eliminate a typicaladditional system on the vehicle stage that performs the settlingfunction.

In addition, another one of the many variations of the present inventioncan be to use a slightly more complex engine cycle than that shown inFIG. 4A or FIG. 4B. Specifically this would involve the use of a kickpump, a secondary pump on the same shaft, on the fuel side of the flowcircuit. Such an arrangement allows for more efficient utilization ofoverall turbopump power, though it does raise the complexity level ofthe configuration.

A person of ordinary skill in the art understand and appreciates thatthe foregoing examples were provided for illustrative purposes and arenot the only ways the present invention can be configured. For example,while FIG. 4A shows the heated fuel drives the turbopumps, there is norequirement of exclusivity. Of course, the introduction of even smallamount of combusted gas to the turbomachinery could create thecatastrophic conditions previously discussed. There could also be morethan heat one exchanger, a plurality of conduits giving and receivingheat, and it is possible that some or all of the fuel flows depictedcould be reversed through the combustion chamber and/or heat exchanger.

1. A gas-generator augmented expander cycle rocket engine, comprising: afirst turbopump for pumping fuel, said first turbopump comprising ashaft having a pump at a first axial end and a turbine at a second axialend; a second turbopump for pumping oxidizer, said first turbopumpcomprising a shaft having a pump at a first axial end and a turbine at asecond axial end; a gas-generator adapted for receiving a first portionof fuel output from said first turbopump and a first portion of oxidizeroutput from said second turbopump and ignite to discharge a heatedcombustion gas; a heat exchanger arranged downstream of saidgas-generator, said heat exchanger being close-coupled to saidgas-generator and having a first conduit for receiving a heated outputof combustion gas from said gas-generator, and a second conduit fortransferring heat arranged adjacent to said first conduit; a thrustchamber assembly comprising a primary combustion chamber at a forwardfacing end and a nozzle at a rearward facing end, said thrust chamberassembly having walls with a plurality of cooling passages therein forreceiving a second portion of fuel output from said first turbopump andtransferring heat to the second portion of the fuel passing through thecooling passages to cool said thrust chamber assembly, wherein saidcooling passages and an input of the second conduit of the heatexchanger are in communication for discharging the second portion offuel from the cooling passages into the input of the second conduit ofsaid heat exchanger to further heat the second portion of fuel; whereinsaid second conduit is in communication with at least one flow path influid communication with a turbine of one or both of said firstturbopump and said second turbopump for discharging said second portionof fuel from said second conduit into the at least one flow path influid communication with a turbine of one or both of said firstturbopump and said second turbopump, and for driving the turbine of oneor both of said first turbopump and said second turbopump by the secondportion of the fuel heated by passage through the second conduit of saidheat exchanger and not exposing said first turbopump and said secondturbopump to combusted gas from either the gas-generator or the primarycombustion chamber; and a fuel output conduit for passing a thirdportion of said fuel output from said first turbopump, and an oxidizeroutput conduit for passing a second portion of oxidizer output from saidsecond turbopump into the primary combustion chamber of said thrustchamber assembly, for forming combustion products for discharge from thenozzle of said thrust chamber assembly to provide thrust to propel therocket.
 2. The rocket engine according to claim 1, wherein said at leastone flow path for receiving the discharge of the second portion of fuelfrom said second conduit comprises a single flow path for exclusivelydriving the turbines of one or both of the first and second turbopumps.3. The rocket engine according to claim 1, wherein the first conduit isin communication with an aft end of the nozzle of the thrust chamberassembly for discharging heated gas from said-gas generator.
 4. Therocket engine according to claim 1, further comprising a secondarynozzle for providing settling thrust, said secondary nozzle beingrelatively smaller than the nozzle of said thrust chamber assembly, andwherein the first conduit is in communication with the secondary nozzlefor introducing combustion gas output from said first conduit to saidsecondary nozzle for providing settling thrust.
 5. The rocket engineaccording to claim 1, wherein said first turbopump further comprises asecondary pump on the same shaft, said secondary pump comprising a kickpump.
 6. The rocket engine according to claim 1, further comprising: agas-generator fuel valve in the flow path of the first portion of fuel,said gas-generator fuel valve for controlling the passage of fuel to thegas-generator; a gas-generator oxidizer valve in the flow path of thefirst portion of oxidizer, said gas-generator oxidizer valve forcontrolling the passage of oxidizer to the gas-generator; a coolantcontrol valve in the flow path of the second portion of fuel, saidcoolant control valve for controlling the passage of coolant to thecooling passages of said thrust chamber assembly; a main fuel value inthe flow path of the third portion of the fuel, said main fuel valve forcontrolling the passage of fuel to the primary combustion chamber ofsaid thrust chamber assembly; and a main oxidizer valve in the flow pathof the second portion of oxidizer, said main oxidizer valve forcontrolling the passage of oxidizer to the primary combustion chamber ofsaid thrust assembly.
 7. The rocket engine according to claim 1, whereinthe third portion of fuel is relatively larger than the first portion offuel and the second portion of fuel, and the second portion of oxidizeris relatively larger than the first portion of oxidizer.
 8. The rocketengine according to claim 6, wherein the gas-generator fuel valve andthe gas-generator oxidizer valve include variable actuators formodulating the output of heated combustion has from said gas-generator.9. The rocket engine according to claim 1, wherein each of the conduitsof said heat exchanger comprises one or more tubes situated cross-wiseto a flow of heated gas from said gas-generator and allowing fuel toflow within and become heated.
 10. The rocket engine according to claim1, wherein the cooling passages in said thrust chamber assemblycommunicate with the second conduit in said heat exchanger and thegas-generator communicates with the first conduit in said heat exchangerfor transferring heat to the second portion of the fuel exiting thecooling passages in said thrust chamber assembly and flowing upstreamthrough the second conduit in said heat exchanger, and for passing aflow of the combustion gas discharged from the gas-generator downstreamthrough the first conduit in said heat exchanger.
 11. The rocket engineaccording to claim 1, wherein the gas-generator is completely decoupledfrom the primary combustion chamber of said thrust chamber assembly. 12.The rocket according to claim 1, wherein said gas-generator includes aninjector at least for purging combustion products from the gas-generatorupon shut down of the rocket engine.
 13. The rocket engine according toclaim 6, further comprising a control module for providing an actuationsequence in said rocket engine, upon a start or restart, in which saidgas-generator fuel valve and the gas-generator oxidizer valve areactuated open prior to opening the coolant control valve, main fuelvalve and main oxidizer valve for igniting the gas-generator prior toigniting said primary combustion chamber for ensuring a smooth enginestart transient.
 14. A gas-generator augmented expander cycle rocketengine comprising: first turbopump means for pumping fuel; secondturbopump means for pumping oxidizer; gas-generator means for combustinga portion of fuel output from said first turbopump means and oxidizerfrom said second turbopump means; a heat exchanger arranged downstreamof said gas-generator means, said heat exchanger being close-coupled tosaid gas-generator means and having a first conduit means for receivinga heated output of gas combusted in said gas-generator means, and asecond conduit means arranged adjacent to said first conduit means fortransferring heat from said first conduit; thrust chamber assembly meanscomprising a primary combustion chamber means facing forward and anozzle means facing rearward, said thrust chamber assembly meansincluding a cooling means milled or channeled in the walls of saidcombustion chamber for passage of a fuel for transferring heat from saidthrust chamber assembly; wherein an output of heated fuel from thecooling means of said thrust chamber assembly means is the sole meansfor driving the first turbopump means and the second turbopump means forthe respective pumping of fuel and oxidizer.
 15. The gas-generatoraugmented expander cycle rocket engine according to claim 14, whereinthe gas combusted in the gas-generator means passes through the heatexchanger means for discharge an aft end of the nozzle means of saidthrust chamber assembly means.
 16. The gas-generator augmented expandercycle rocket engine according to claim 14, further comprising asecondary nozzle means for providing settling thrust, wherein the gascombusted in the gas-generator is discharged through the secondarynozzle means so as to provide settling thrust, and wherein saidsecondary nozzle means is relatively smaller than the nozzle means ofsaid thrust chamber assembly means.
 17. A method for operating agas-generator augmented expander cycle rocket engine, comprising thesteps of: providing a first turbopump for pumping fuel, and a secondturbopump means for pumping oxidizer; combusting a portion of fueloutput from said first turbopump means and oxidizer from said secondturbopump means in a gas-generator; arranging a heat exchangerdownstream of said gas-generator that is close-coupled to saidgas-generator, outputting a heated output of combusted gas from saidgas-generator into a first conduit of said heat exchanger, combusting amain portion of fuel and oxidizer in a primary combustion chamber of athrust chamber assembly that is decoupled from said gas generator saidthrust chamber assembly being provided with cooling passages milled orchanneled in the walls of said combustion chamber; passing a thirdportion of fuel through the cooling passages of said combustion chamberto transfer heat from said thrust chamber assembly to form heated fuel;discharging the heated fuel into a second conduit of said heatexchanger; and discharging the fuel from the second conduit of said heatexchanger into a flow path of the turbines of the first and secondturbopumps so that turbines of the turbopumps are exclusively driven bythe heated fuel and the turbopumps are not exposed to combustion gasfrom the primary combustion chamber or the gas-generator.